This invention relates to a cooled turbine aerofoil for use in a gas turbine engine.
The second law of thermodynamics dictates that the thermal efficiency of the jet engine cycle is increased by increasing the turbine entry temperature. Turbine entry temperatures of modern gas turbine engines are in excess of the melting point temperatures of the turbine components and therefore some form of cooling of the turbine aerofoils is required, that is to say the nozzle guide vane and turbine rotor blade aerofoils. Typically the aerofoils are provided with internal passages or cavities through which cooling air bleed from the engine compressor is directed. In the case of turbine rotor blades the cooling air is generally fed into inlets provided in the root portion of the blade. The cooling passages or cavities direct the cooling air through single or multi pass channels with some of the cooling air being exhausted from the internal passages through film cooling holes provided in the aerofoil external surfaces. Film cooling holes provide additional cooling of these external aerofoil surfaces where the exiting cooling air provides a thin film of relatively cool air over the aerofoil surface in the region of the film-cooling hole to protect the aerofoil from the high temperature turbine gases.
As turbine entry temperatures have increased it has been necessary to use greater amounts of cooling air from the engine compressor. However, there is a limit to the amount of cooling air that can be used for turbine cooling since increasing the amount of cooling air bled from the compressor results in reduced engine cycle efficiency and can lead to increased smoke and/or noxious gaseous emissions. Hitherto, it has be possible to increase the cooling effectiveness of the cooling air in order to minimise the amount of cooling air that is bled from the compressor.
The leading edge of air cooled aerofoils is usually provided with one or more rows of film cooling holes in the vicinity of the leading edge stagnation point, that is to say in the point on the aerofoil cross-section with the highest external pressure. Cooling the leading edge of a turbine aerofoil therefore requires that the cooling air is supplied at a pressure higher than that at the stagnation point. This is necessary to ensure that the cooling air can flow from within the respective aerofoil internal passage or cavity to the external surfaces without the possibility of the high temperature engine gases being ingested into the film cooling holes. Supplying cooling air at such high pressure has a number of consequences, in particular increased combustor pressures and reduced thermodynamic cycle deficiency.